1. Field of Endeavor
The present invention relates to gas turbines, and more specifically to an axial flow gas turbine.
Yet more specifically, the invention relates to designing a stage of an axial flow turbine for a gas turbine unit. Generally the turbine stator has a vane carrier with slots where a row of vanes and a row of stator heat shields are installed one after another. The same stage includes a rotor with a rotating shaft with slots where a row of rotor heat shields and a row of blades are installed one after another.
2. Brief Description of the Related Art
The invention relates to a gas turbine of the axial flow type, an example of which is shown in FIG. 1. The gas turbine 10 of FIG. 1 operates according to the principle of sequential combustion. It includes a compressor 11, a first combustion chamber 14 with a plurality of burners 13 and a first fuel supply 12, a high-pressure turbine 15, a second combustion chamber 17 with the second fuel supply 16, and a low-pressure turbine 18 with alternating rows of blades 20 and vanes 21, which are arranged in a plurality of turbine stages arranged along the machine axis MA.
The gas turbine 10 according to FIG. 1 includes a stator and a rotor. The stator includes a vane carrier 19 with the vanes 21 mounted therein; these vanes 21 are necessary to form profiled channels where hot gas developed in the combustion chamber 17 flows through. Gas flowing through the hot gas path 22 in the required direction hits against the blades 20 installed in shaft slits of a rotor shaft and causes the turbine rotor to rotate. To protect the stator housing against the hot gas flowing above the blades 20, stator heat shields installed between adjacent vane rows are used. High temperature turbine stages require cooling air to be supplied into vanes, stator heat shields and blades.
To ensure high-temperature turbine stage operation with a long-term lifespan, all parts of the hot gas path 22 should be cooled effectively. Parts of the known design presented in FIGS. 2(a) and (b) are cooled as follows. Compressed cooling air 24 is delivered from the compressor through a plenum 23 and enters cavities 31 and 29. In the case of cavity 31, this is done by a hole 25. Then this cooling air flows out from the airfoil of vane 21 and through holes 30 and 28 of the stator heat shield 27, which is attached to an inner ring 26 opposite to the blade 20, into the turbine flow path 22. The thin-walled crown 32 (FIG. 2(b)) of the peripheral blade zone (the blade tip) is very sensitive to high gas temperature. Cooling air escaping from holes 30 situated in the forward part of the stator heat shield 27 in the design of FIG. 2 contributes to lowering the temperature of the blade crown 32 (in addition to the blade cooling system itself, which is not explicitly shown in the figure).
However, the above described design can have the following disadvantages:
1. Due to the large distance from the outlets of holes 30 to the leading edge of the blades 20, cooling air jets soon lose their energy and are washed out with hot gas from the hot gas path 22.
2. Air flowing out of holes 30 has a rather high temperature, since it has already cooled a substantial surface area of the stator heat shield 27.
3. No effective blowing through with cooling air is provided for the space between adjacent stator heat shields 27 (FIG. 2(b)), and this increases the overheating risk for sealing plates 33 and side surfaces of the stator heat shields 27.